1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled turbine rotor blade with both near wall cooling of the airfoil and sealing of the blade tip.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
For a blade cooled with radial flow channels formed within the walls, the near wall radial flow channel at the blade tip discharge section experiences an external cross flow effect. As a result of this cross flow effect, an over-temperature occurs at the locations of the blade tip on the pressure wall side. This external cross flow effect on the near wall radial flow channel is caused by a non-uniformity of the radial channel discharge pressure profile and the blade tip leakage flow across the radial channel exit location.
One process for cooling a turbine rotor blade is disclosed in U.S. Pat. No. 5,702,232 issued to Moore on Dec. 30, 1997 and entitled COOLED AIRFOILS FOR A GAS TURBINE ENGINE. In the Moore blade cooling design, the blade mid-chord section is cooled with a number of radial extending single pass cooling channels that open onto the blade tip. A radial cooling channel can be of a race-track shape instead of circular. Film cooling holes are also connected to the radial cooling channels to discharge layers of film cooling air onto the external blade surface. In this design, cooling flow velocity decreases with passage through the channel and thus the internal heat transfer coefficient is reduced. Cooling air refresh holes are therefore used that bring cooling air from a central cavity and into the radial cooling channels to replenish the cooling air flow.